This application relates generally to gas turbine engines and, more particularly, to methods and apparatus for cooling airfoils used within gas turbine engines.
At least some known gas turbine engines include a compressor, a combustor, and a turbine. Airflow entering the compressor is compressed and directed to the combustor where it is mixed with fuel and ignited, producing hot combustion gases used to drive the turbine. Because components within the turbine are exposed to hot combustion gases, cooling air is routed to the airfoils and blades.
For example, a turbine vane or rotor blade typically includes a hollow airfoil, the outside of which is exposed to the hot combustion gases, and the inside of which is supplied with cooling fluid, which is typically compressed air. The airfoil includes leading and trailing edges, a pressure side, and a suction side. The pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between an airfoil root and an airfoil tip. Film cooling holes extend between a cooling chamber defined within the airfoil and an outer surface of the airfoil. The cooling holes route cooling fluid from the cooling chamber to the outside of the airfoil for film cooling the airfoil. The film cooling holes discharge cooling fluid at an injection angle that is measured with respect to the outer surface of the airfoil.
Because of the curvature distribution of the outer surface of the airfoil between the leading and trailing edges, the injection angles of the cooling holes are typically between 25 and 40 degrees. Cooling fluid discharged from cooling holes having increased injection angles may separate from the surface of the airfoil and mix with the hot combustion gases. Such separation decreases an effectiveness of the film cooling and increases aerodynamic mixing losses.
To facilitate reducing aerodynamic mixing losses, at least some known airfoils include curved film cooling openings. The curved film cooling openings have injection angles as low as 16.5 degrees. However, the cooling fluid may separate from an inner wall of the cooling opening and be discharged in an erratic manner. Furthermore, manufacturing such curved openings is a complex and costly procedure.
In an exemplary embodiment, an airfoil for a gas turbine engine includes an inflection that facilitates enhancing film cooling of the airfoil, without adversely impacting aerodynamic efficiency of the airfoil. The airfoil includes a generally concave first sidewall and a generally convex second sidewall. The sidewalls are joined at a leading edge and at a chordwise spaced trailing edge of the airfoil that is downstream from leading edge. A cooling chamber is defined within the sidewalls, and a plurality of cooling openings extend between the cooling chamber and an external surface of the first sidewall. At least one of the cooling openings extends from the cooling chamber into the inflection at an injection angle measured with respect to an external surface of the airfoil.
In another aspect, a gas turbine engine including a plurality of airfoils that each include a leading edge, a trailing edge, a first sidewall having an outer surface, and a second sidewall having an outer surface is provided. The airfoil first and second sidewalls are connected chordwise at the leading and trailing edges. The first and second sidewalls extend radially from an airfoil root to an airfoil tip, and at least one of the first sidewall and said second sidewall also includes an inflection.
In a further aspect, a method for contouring an airfoil for a gas turbine engine to facilitate improving film cooling effectiveness of the airfoil is provided. The airfoil includes a leading edge, a trailing edge, a first sidewall, and a second sidewall. The first and second sidewalls are connected chordwise at the leading and trailing edges to define a cavity, and extend radially between an airfoil root and an airfoil tip. The method includes the steps of forming an inflection in an outer surface of at least one of the airfoil first sidewall and the airfoil second sidewall, such that the inflection extends a distance radially between the airfoil root and the airfoil tip, and forming at least one opening within the inflection for receiving cooling fluid therethrough from the airfoil cavity to the airfoil outer surface.